Referring to FIG. 1, a gas turbine engine is generally indicated at 10 and comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high pressure compressor 14, a combustor 15, a turbine arrangement comprising a high pressure turbine 16, an intermediate pressure turbine 17 and a low pressure turbine 18, and an exhaust nozzle 19.
The thermodynamic efficiency (i.e. the fuel burn at a given thrust) is determined in part by the temperature of the working fluids entering the high, intermediate and low pressure turbines 16-18 (known as turbine entry temperature, TET). The TET in modern turbines often approaches or exceeds the melting point of the materials used to make the turbine 17, 18, 19. Consequently, cooling air is provided to cool one or more components (such as blades, stators and discs) of one or more turbine stages 16-18 to prevent damage to the components that would otherwise be caused by the high temperature working fluids, or to increase component life.
However, increasing the cooling air mass flow reduces the thermodynamic efficiency of the engine, since the air in the cooling flow is no longer available to do useful work in the thermodynamic cycle. In many cases, the required cooling air mass flow may be relatively high, since the temperature of the compressed cooling air may be high (typically up to around 700° C.).
One proposed solution to help improve cooling efficiency is to cool the cooling air to a lower temperature prior to delivery. Such systems are known as “cooled cooling air” systems.
FIG. 2 shows a schematic representation of a portion of a gas turbine engine 10 centred around the combustor which includes a cooled cooling air system 20. In the system 20, relatively low temperature air provided by the propulsive fan 12 flows through an air to air heat exchanger 22 along a first fluid path 24. Compressed cooling air from the high pressure compressor 14 flows through the heat exchanger 22 along a second fluid path 26 which is in thermal contact with the first fluid path 24. Consequently, the cooling air in the second fluid path 26 is cooled prior to delivery to the components which require cooling.
In current systems, the flow of fan air through the first path 24 is only activated in accordance with one or more predetermined conditions. In one known system, an Electronic Engine Control Unit (EEC) monitors the high pressure compressor delivery temperature, i.e. the temperature of the air in the second fluid path 26 downstream of the heat exchanger. When T30 is above a predetermined value, a valve 30 is actuated such that the fan air flows through the first path 24 to provide the required cooling. The first path 24 may also be used to determine the temperature of the cooling air supply which can be used to control the required flow.
Other types of cooled cooling air systems and concepts are known, such as fuel cooled cooling air, in which engine fuel is used to cool the cooling fluid. Such systems may be operated in a similar manner.
Another important role for cooling systems is to forceably cool the engine after a period of operation in order to help extend an acceptable service life between replacement or overhaul. This requires the engine to be run at low power to provide a “cool down” period, particularly after extended use at high thrust. One way to achieve this is to mandate a running period after landing to provide an air flow through the engine for a sufficient period in order to cool various components. However, this particular approach burns excessive fuel and engines are often shutdown prior to the full cool down period in practice in order to save fuel. This results in deterioration of engine components and a reduced service life of an engine. In either case, increased costs result due to excess fuel burn or reduced service life.
The present invention provides a method of controlling an aircraft gas turbine engine which seeks to overcome some or all of the above problems.